Platform cooling circuit for a gas turbine engine component

ABSTRACT

A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform having a first path side and a second path side and a platform cooling circuit disposed on one of the first path side and the second path side of the platform. The platform cooling circuit includes a first core cavity, a cavity in fluid communication with the first core cavity, and a cover plate positioned to cover at least the cavity.

CROSS REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. patent application Ser. No.13/586,110, which was filed on Aug. 15, 2012.

BACKGROUND

This disclosure relates generally to a gas turbine engine, and moreparticularly to a component that can be incorporated into a gas turbineengine. The component can include a platform cooling circuit for coolinga platform of the component.

Gas turbine engines typically include a compressor section, a combustorsection and a turbine section. During operation, air is pressurized inthe compressor section and is mixed with fuel and burned in thecombustor section to generate hot combustion gases. The hot combustiongases are communicated through the turbine section, which extractsenergy from the hot combustion gases to power the compressor section andother gas turbine engine loads.

Both the compressor and turbine sections of the gas turbine engine mayinclude alternating rows of rotating blades and stationary vanes thatextend into the core flow path of the gas turbine engine. For example,in the turbine section, turbine blades rotate to extract energy from thehot combustion gases that are communicated along the core flow path ofthe gas turbine engine. The turbine vanes prepare the airflow for thenext set of blades. These blades and vanes are examples of componentsthat may need cooled by a dedicated source of cooling airflow in orderto withstand the relatively high temperatures of the hot combustiongases that are communicated along the core flow path.

SUMMARY

A component for a gas turbine engine according to an exemplary aspect ofthe present disclosure includes, among other things, a platform having afirst path side and a second path side and a platform cooling circuitdisposed on one of the first path side and the second path side of theplatform. The platform cooling circuit includes a first core cavity, acavity in fluid communication with the first core cavity, and a coverplate positioned to cover at least the cavity.

In a further non-limiting embodiment of the foregoing component, thefirst path side is a non-gas path side and the second path side is a gaspath side.

In a further non-limiting embodiment of either of the forgoingcomponents, the component is a turbine vane.

In a further non-limiting embodiment of any of the foregoing components,the cavity is a serpentine cavity.

In a further non-limiting embodiment of any of the foregoing components,the platform cooling circuit includes an impingement cavity that isseparate from the cavity.

In a further non-limiting embodiment of any of the foregoing components,the cover plate is positioned to cover the impingement cavity.

In a further non-limiting embodiment of any of the foregoing components,the cavity includes a first serpentine portion in fluid communicationwith the first core cavity and a second serpentine portion in fluidcommunication with a second core cavity.

In a further non-limiting embodiment of any of the foregoing components,a communication passage connects the first serpentine portion to thesecond serpentine portion.

In a further non-limiting embodiment of any of the foregoing components,the cavity includes a plurality of augmentation features.

In a further non-limiting embodiment of any of the foregoing components,there are a plurality of openings through the cover plate.

In a further non-limiting embodiment of any of the foregoing components,a second core cavity is in fluid communication with the cavity. Thesecond core cavity is positioned adjacent to the first core cavity onthe platform.

A gas turbine engine according to an exemplary aspect of the presentdisclosure includes, among other things, a compressor section and acombustor section in fluid communication with the compressor section. Aturbine section is in fluid communication with the combustor section.One of the compressor section and the turbine section includes at leastone component having a platform that includes a platform coolingcircuit. The platform cooling circuit includes a first core cavity, acavity in fluid communication with the first core cavity and a coverplate positioned to cover at least the cavity.

In a further non-limiting embodiment of the foregoing gas turbineengine, the at least one component is a vane.

In a further non-limiting embodiment of either of the foregoing gasturbine engines, the cavity is a serpentine cavity.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, a second core cavity is in fluid communication with the cavityand positioned adjacent to the first core cavity on the platform.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the platform cooling circuit includes an impingement cavitythat is separate from the cavity.

A method of cooling a component of a gas turbine engine according toanother exemplary aspect of the present disclosure includes, among otherthings, communicating a cooling airflow into a first core cavity of aplatform of the component and circulating the cooling airflow from thefirst core cavity through a first portion of a cavity that is in fluidcommunication with the first core cavity.

In a further non-limiting embodiment of the foregoing method of coolinga component of a gas turbine engine, the cooling airflow can becommunicated into a second core cavity of the platform and circulatedfrom the second core cavity through a second portion of the cavity thatis in fluid communication with the second core cavity.

In a further non-limiting embodiment of either of the foregoing methodsof cooling a component of a gas turbine engine, the cavity is aserpentine cavity.

In a further non-limiting embodiment of any of the foregoing methods ofcooling a component of a gas turbine engine, the step of communicatingincludes communicating the cooling airflow through at least one feedopening in a leading edge portion of the platform and into the firstcore cavity.

The various features and advantages of this disclosure will becomeapparent to those skilled in the art from the following detaileddescription. The drawings that accompany the detailed description can bebriefly described as follows.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 illustrates a schematic, cross-sectional view of a gas turbineengine.

FIG. 2 illustrates a component that can be incorporated into a gasturbine engine.

FIG. 3 illustrates a platform of a component that can be incorporatedinto a gas turbine engine.

FIG. 4 illustrates an exemplary platform cooling circuit of a platformof a component.

FIG. 5 illustrates another exemplary platform cooling circuit.

FIG. 6 illustrates yet another exemplary platform cooling circuit.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The exemplarygas turbine engine 20 is a two-spool turbofan engine that generallyincorporates a fan section 22, a compressor section 24, a combustorsection 26 and a turbine section 28. Alternative engines might includean augmenter section (not shown) among other systems for features. Thefan section 22 drives air along a bypass flow path B, while thecompressor section 24 drives air along a core flow path C forcompression and communication into the combustor section 26. The hotcombustion gases generated in the combustor section 26 are expandedthrough the turbine section 28. Although depicted as a turbofan gasturbine engine in the disclosed non-limiting embodiment, it should beunderstood that the concepts described herein are not limited toturbofan engines and these teachings could extend to other types ofengines, including but not limited to, three-spool engine architectures.

The gas turbine engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centerlinelongitudinal axis A. The low speed spool 30 and the high speed spool 32may be mounted relative to an engine static structure 33 via severalbearing systems 31. It should be understood that other bearing systems31 may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 34 thatinterconnects a fan 36, a low pressure compressor 38 and a low pressureturbine 39. The inner shaft 34 can be connected to the fan 36 through ageared architecture 45 to drive the fan 36 at a lower speed than the lowspeed spool 30. The high speed spool 32 includes an outer shaft 35 thatinterconnects a high pressure compressor 37 and a high pressure turbine40. In this embodiment, the inner shaft 34 and the outer shaft 35 aresupported at various axial locations by bearing systems 31 positionedwithin the engine static structure 33.

A combustor 42 is arranged between the high pressure compressor 37 andthe high pressure turbine 40. A mid-turbine frame 44 may be arrangedgenerally between the high pressure turbine 40 and the low pressureturbine 39. The mid-turbine frame 44 can support one or more bearingsystems 31 of the turbine section 28. The mid-turbine frame 44 mayinclude one or more airfoils 46 that extend within the core flow path C.

The inner shaft 34 and the outer shaft 35 are concentric and rotate viathe bearing systems 31 about the engine centerline longitudinal axis A,which is co-linear with their longitudinal axes. The core airflow iscompressed by the low pressure compressor 38 and the high pressurecompressor 37, is mixed with fuel and burned in the combustor 42, and isthen expanded over the high pressure turbine 40 and the low pressureturbine 39. The high pressure turbine 40 and the low pressure turbine 39rotationally drive the respective high speed spool 32 and the low speedspool 30 in response to the expansion.

In a non-limiting embodiment, the gas turbine engine 20 is a high-bypassgeared aircraft engine. In a further example, the gas turbine engine 20bypass ratio is greater than about six (6:1). The geared architecture 45can include an epicyclic gear train, such as a planetary gear system orother gear system. The example epicyclic gear train has a gear reductionratio of greater than about 2.3, and in another example is greater thanabout 2.5:1. The geared turbofan enables operation of the low speedspool 30 at higher speeds, which can increase the operational efficiencyof the low pressure compressor 38 and low pressure turbine 39 and renderincreased pressure in a fewer number of stages.

The pressure ratio of the low pressure turbine 39 can be pressuremeasured prior to the inlet of the low pressure turbine 39 as related tothe pressure at the outlet of the low pressure turbine 39 and prior toan exhaust nozzle of the gas turbine engine 20. In one non-limitingembodiment, the bypass ratio of the gas turbine engine 20 is greaterthan about ten (10:1), the fan diameter is significantly larger thanthat of the low pressure compressor 38, and the low pressure turbine 39has a pressure ratio that is greater than about 5 (5:1). It should beunderstood, however, that the above parameters are only exemplary of oneembodiment of a geared architecture engine and that the presentdisclosure is applicable to other gas turbine engines, including directdrive turbofans.

In this embodiment of the exemplary gas turbine engine 20, a significantamount of thrust is provided by the bypass flow path B due to the highbypass ratio. The fan section 22 of the gas turbine engine 20 isdesigned for a particular flight condition—typically cruise at about 0.8Mach and about 35,000 feet. This flight condition, with the gas turbineengine 20 at its best fuel consumption, is also known as bucket cruiseThrust Specific Fuel Consumption (TSFC). TSFC is an industry standardparameter of fuel consumption per unit of thrust.

Fan Pressure Ratio is the pressure ratio across a blade of the fansection 22 without the use of a Fan Exit Guide Vane system. The low FanPressure Ratio according to one non-limiting embodiment of the examplegas turbine engine 20 is less than 1.45. Low Corrected Fan Tip Speed isthe actual fan tip speed divided by an industry standard temperaturecorrection of “T”/518.7^(0.5), where T represents the ambienttemperature in degrees Rankine. The Low Corrected Fan Tip Speedaccording to one non-limiting embodiment of the example gas turbineengine 20 is less than about 1150 fps (351 m/s). The teachings of thisdisclosure could also extend to non-geared, low-bypass, no-bypass andindustrial applications.

Each of the compressor section 24 and the turbine section 28 may includealternating rows of rotor assemblies and vane assemblies (shownschematically) that carry airfoils that extend into the core flow pathC. For example, the rotor assemblies can carry a plurality of rotatingblades 25, while each vane assembly can carry a plurality of vanes 27that extend into the core flow path C. The blades 25 of the rotorassemblies create or extract energy (in the form of pressure) from thecore airflow that is communicated through the gas turbine engine 20along the core flow path C. The vanes 27 of the vane assemblies directthe core airflow to the blades 25 to either add or extract energy.

Various components of a gas turbine engine 20, such as the airfoils ofthe blades 25 and the vanes 27 of the compressor section 24 and theturbine section 28, may be subjected to repetitive thermal cycling underwidely ranging temperatures and pressures. The hardware of the turbinesection 28 is particularly subjected to relatively extreme operatingconditions. Therefore, some components may require cooling circuits forcooling the parts during engine operation. Example platform coolingcircuits for cooling a platform of a component are discussed below.

FIG. 2 illustrates a component 50 that can be incorporated into a gasturbine engine, such as the gas turbine engine 20 of FIG. 1. In thisembodiment, the component 50 is a turbine vane. However, the teachingsof this disclosure are not limited to turbine vanes and could extend toother components of the gas turbine engine 20, including but not limitedto, compressor blades and vanes, turbine blades, blade outer air seals(BOAS), mid-turbine frames, transition ducts, or any other componentthat extends within the core flow path C.

The component 50 can include one or more platforms 52 and one or moreairfoils 54 that extend from the platform(s) 52. In this particularembodiment, the component 50 includes an inner diameter platform 52A andan outer diameter platform 52B as well as two airfoils 54A, 54B thatextend between the inner and outer platforms 52A, 52B. Althoughillustrated as a vane doublet, it should be understood that vanesinglets or other vane assemblies may benefit from the teachings of thisdisclosure, and that non-airfoil components, such as BOAS or transitionducts, may also benefit from these teachings.

The platform(s) 52 include a leading edge portion 56, a trailing edgeportion 58, and opposing mate faces 60, 62. The platform(s) 52 axiallyextend between the leading edge portion 56 and the trailing edge portion58 and circumferentially extend between the opposing mate faces 60, 62.The opposing mate faces 60, 62 can be mounted relative to correspondingmate faces of adjacent components of a gas turbine engine to provide afull ring assembly, such as a full ring vane assembly that can becircumferentially disposed about the engine centerline longitudinal axisA (see FIG. 1).

The platforms can also include a first path side (for example, a non-gaspath side) 64 and a second path side (for example a gas path side) 66.In other words, when the component 50 is mounted within the gas turbineengine 20, the non-gas path side 64 is positioned on a non-core flowpath side of the component 50, while the gas path side 66 may establishan outer boundary of the core flow path C of the gas turbine engine 20.

One or both of the platforms 52 can also include a platform coolingcircuit 68 for cooling the platform 52. A cooling airflow 70 can becirculated through the platform cooling circuit 68 to transfer thermalenergy from the platform 52 to the cooling airflow 70, thereby coolingportions of the platform 52. The cooling airflow 70 is communicated tothe platform cooling circuit 68 from an airflow source 72 that isexternal to the component 50. The cooling airflow 70 is of a lowertemperature than the airflow communicated along the core flow path Cthat is communicated across the component 50. In one embodiment, thecooling airflow 70 is a bleed airflow that can be sourced from thecompressor section 24 or any other portion of the gas turbine enginethat is upstream from the component 50. The cooling airflow 70 couldalso be a recycled cooling airflow that is first used to cool anupstream component of the gas turbine engine 20, such as an upstreamBOAS, for example.

One exemplary platform cooling circuit 68 is illustrated by FIG. 3. Theplatform cooling circuit 68 is disposed on the non-gas path side 64 ofthe platform 52. In this embodiment, the platform 52 could berepresentative of either an inner diameter platform or an outer diameterplatform of a vane, or could be a platform of some other component,including but not limited to, a blade or a BOAS.

The exemplary platform cooling circuit 68 can include a first corecavity 74, a second core cavity 76 that is adjacent to the first corecavity 74, a serpentine cavity 78 (i.e., a third cavity) that is influid communication with each of the first core cavity 74 and the secondcore cavity 76, and a cover plate 80 positioned at the non-gas path side64 of the platform 52. It should be understood that the platform coolingcircuit 68 could include fewer or additional core cavities and is notnecessarily limited to the particular configuration shown in FIG. 3.

The first core cavity 74 and the second core cavity 76 can receivecooling airflow 70 from the airflow source 72 to provide convectivecooling to the leading edge portion 56 of the platform 52 and theserpentine cavity 78. In one non-limiting embodiment, the first corecavity 74 and the second core cavity 76 are produced using a ceramiccore and the serpentine cavity 78 is produced using a wax form or aceramic core.

In this embodiment, the cover plate 80 is positioned to cover theserpentine cavity 78 to define an enclosed cooling passage therein. Thecover plate 80 may be brazed or welded onto the non-gas path side 64 ofthe platform 52. The cover plate 80 may be shaped similar to theserpentine cavity 78.

The platform cooling circuit 68 can further include an impingementcavity 82 that is separate from the serpentine cavity 78. Theimpingement cavity 82 is positioned adjacent to the trailing edgeportion 58 of the platform 52 and may extend from a position adjacent tothe mate face 60 toward the opposing mate face 62. The impingementcavity 82 of this embodiment is generally L-shaped, although othershapes are also contemplated. The cover plate 80 may extend to alsocover the impingement cavity 82.

In a non-limiting embodiment, the cover plate 80 may include one or moreopenings 84 that extend through the cover plate 80. Impingement coolingairflow 86 can be communicated through the openings 84 into portions ofone or both of the serpentine cavity 78 and the impingement cavity 82 toimpingement cool the surfaces of these cavities. The impingement coolingairflow 86 can be sourced from the airflow source 72 or from a separatesource.

FIG. 4 illustrates the platform 52 with the cover plate 80 removed tobetter illustrate several features of the platform cooling circuit 68.In this embodiment, the serpentine cavity 78 includes a first serpentineportion 88 and a second serpentine portion 90. A communication passage92 can connect the first serpentine portion 88 to the second serpentineportion 90. In one embodiment, a portion of the openings 84 of the coverplate 80 are positioned to direct the impingement cooling airflow 86into the communication passage 92 (see FIG. 3). The communicationpassage 92 can be a cast or machined feature of the platform coolingcircuit 68.

The first serpentine portion 88 of the serpentine cavity 78 is in fluidcommunication with the first core cavity 74, and the second serpentineportion 90 is in fluid communication with the second core cavity 76. Thefirst core cavity 74 may feed into an inlet 94 of the first serpentineportion 88, and the second core cavity 76 may feed into an inlet 96 ofthe second serpentine portion 90.

FIG. 5 schematically illustrates a method of cooling the component 50using the exemplary platform cooling circuit 68. Cooling airflow 70 canbe communicated into each of the first core cavity 74 and the secondcore cavity 76. The cooling airflow 70 is communicated through one ormore feed openings 98 that can extend through the leading edge portion56 of the platform 52. In one embodiment, the feed openings 98 extendthrough an upstream wall 100 of the leading edge portion 56 (see FIG.2). However, the feed openings 98 can be positioned at any location ofthe platform 52 to direct the cooling airflow 70 into the first corecavity 74 and the second core cavity 76.

The cooling airflow 70 that is communicated into the first core cavity74 can be directed to the inlet 94 of the first serpentine portion 88 ofthe serpentine cavity 78. The cooling airflow 70 can then becommunicated in a serpentine path that is established by the serpentinecavity 78 to cool the platform 52. For example, the cooling airflow 70can first be directed from the inlet 94 in a first direction D1 towardthe trailing edge portion 58 within a first passage 89 of the firstserpentine portion 88. The cooling airflow 70 is then communicatedthrough a turn 91 of the first serpentine portion 88 prior to entering asecond passage 93. The cooling airflow 70 is communicated within thesecond passage 93 in a second direction D2 that is opposite of the firstdirection D1. The cooling airflow 70 can exit the first serpentineportion 88 through film openings 95 disposed within the second passage93. For example, the cooling airflow 70 can be returned to the core flowpath C.

Meanwhile, the cooling airflow 70 from the second core cavity 76 can becommunicated through the inlet 96 of the second serpentine portion 90 ofthe serpentine cavity 78 and into a first passage 97. The coolingairflow 70 is directed in the first direction D1 within the firstpassage 97 before entering a turn 99. From the turn 99, the coolingairflow 70 can be communicated in the second direction D2 within asecond passage 101. In other words, the cooling airflow 70 is circulatedthrough the second serpentine portion 90 to cool the platform 52. Thecooling airflow 70 can exit the second serpentine portion 90 throughfilm openings 95 disposed within the second passage 101.

Portions of the cooling airflows 70 of both the first serpentine portion88 and the second serpentine portion 90 may be intermixed within thecommunication passage 92. In this embodiment, the communication passage92 extends between the second passage 93 of the first serpentine portion88 and the first passage 97 of the second serpentine portion 90.

The platform cooling circuit 68 can also impingement cool portions ofthe platform 52, such as described above via the openings 84 of thecover plate 80 (see FIG. 3). Portions P1 of the cooling airflow 70 mayalso be communicated into a cavity 102 of the component 50 to cool otherportions of the component 50, such as an airfoil portion. The portionsP1 of the cooling airflow 70 can enter the cavity 102 through one orboth of the first core cavity 74 and the second core cavity 76.

FIG. 6 illustrates another exemplary platform cooling circuit 168 of aplatform 152 of a component 150. The platform cooling circuit 168 issimilar to the platform cooling circuit 68 described above but mayinclude slightly modified features. In this disclosure, like referencenumerals signify like or similar features, whereas reference numeralsmodified by “100” signify features that have been modified.

In this embodiment, the platform cooling circuit 168 includes aserpentine cavity 178 having a plurality of augmentation features 104.The plurality of augmentation features 104 may include any heat transferaugmentation feature such as pins, linear trip strips, or chevron tripstrips. In another embodiment, the impingement cavity 82 also includesaugmentation features 104.

Although the different non-limiting embodiments are illustrated ashaving specific components, the embodiments of this disclosure are notlimited to those particular combinations. It is possible to use some ofthe components or features from any of the non-limiting embodiments incombination with features or components from any of the othernon-limiting embodiments.

It should be understood that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be understood that although a particular componentarrangement is disclosed and illustrated in these exemplary embodiments,other arrangements could also benefit from the teachings of thisdisclosure.

The foregoing description shall be interpreted as illustrative and notin any limiting sense. A worker of ordinary skill in the art wouldrecognize that various modifications could come within the scope of thisdisclosure. For these reasons, the following claims should be studied todetermine the true scope and content of this disclosure.

What is claimed is:
 1. A component for a gas turbine engine, comprising:a platform having a first path side and a second path side; and aplatform cooling circuit disposed on one of said first path side andsaid second path side of said platform, wherein said platform coolingcircuit includes: a first core cavity; a serpentine cavity in fluidcommunication with said first core cavity; a cover plate shaped similarto the serpentine cavity and positioned to cover at least saidserpentine cavity, wherein said first core cavity is at least partiallyuncovered by said cover plate; a plurality of openings through saidcover plate, wherein said openings are configured to communicateimpingement cooling airflow to said serpentine cavity.
 2. The componentas recited in claim 1, wherein said first path side is a non-gas pathside and said second path side is a gas path side.
 3. The component asrecited in claim 1, wherein the component is a turbine vane.
 4. Thecomponent as recited in claim 1, wherein said platform cooling circuitincludes an impingement cavity that is separate from said serpentinecavity.
 5. The component as recited in claim 4, wherein said cover plateis positioned to cover said impingement cavity.
 6. The component asrecited in claim 1, wherein said serpentine cavity includes a firstserpentine portion in fluid communication with said first core cavityand a second serpentine portion in fluid communication with a secondcore cavity, and comprising a communication passage that connects saidfirst serpentine portion to said second serpentine portion, wherein theplurality of openings are positioned to direct impingement airflow intothe communication passage.
 7. The component as recited in claim 1,wherein said serpentine cavity includes a plurality of augmentationfeatures.
 8. The component as recited in claim 1, comprising a secondcore cavity that is in fluid communication with said serpentine cavity,wherein said second core cavity is positioned adjacent to said firstcore cavity on said platform.
 9. The component as recited in claim 1,wherein said platform includes a feed opening at an edge of saidplatform, and said feed opening is in direct fluid communication withsaid first core cavity.
 10. A component for a gas turbine engine,comprising: a platform having a first path side and a second path side;and a platform cooling circuit disposed on one of said first path sideand said second path side of said platform, wherein said platformcooling circuit includes: a first core cavity; a serpentine cavity influid communication with said first core cavity; a cover plate shapedsimilar to the serpentine cavity and positioned to cover at least saidserpentine cavity, wherein said first core cavity is at least partiallyuncovered by said cover plate, wherein said platform cooling circuitincludes a second core cavity adjacent said first core cavity and influid communication with said serpentine cavity, wherein said secondcore cavity is at least partially uncovered by said cover plate.
 11. Acomponent for a gas turbine engine, comprising: a platform having afirst path side and a second path side; and a platform cooling circuitdisposed on one of said first path side and said second path side ofsaid platform, wherein said platform cooling circuit includes: a firstcore cavity; a serpentine cavity in fluid communication with said firstcore cavity; an impingement cavity that is separate from said serpentinecavity; and a cover plate shaped similar to said serpentine cavity andpositioned to cover at least said serpentine cavity and said impingementcavity, wherein said impingement cavity is generally L-shaped, and theplatform extends circumferentially between first and second mate faces,and the impingement cavity is positioned adjacent a trailing edge of theplatform and extends from a position adjacent the first mate face to aposition adjacent the second mate face.
 12. The component as recited inclaim 11, comprising a plurality of openings through said cover plate,wherein said openings are configured to communicate impingement coolingairflow into said impingement cavity.
 13. The component as recited inclaim 11, comprising a plurality of openings through said cover plate,wherein said openings are configured to communicate impingement coolingairflow into said serpentine cavity.
 14. The component as recited inclaim 11, wherein said platform cooling circuit includes a second corecavity adjacent said first core cavity and in fluid communication withsaid serpentine cavity.
 15. The component as recited in claim 11,wherein the impingement cavity includes augmentation features.
 16. Acomponent for a gas turbine engine, comprising: a platform having afirst path side and a second path side; and a platform cooling circuitdisposed on one of said first path side and said second path side ofsaid platform, wherein said platform cooling circuit includes: a firstcore cavity; a serpentine cavity in fluid communication with said firstcore cavity; a cover plate shaped similar to the serpentine cavity andpositioned to cover at least said serpentine cavity, wherein said firstcore cavity is at least partially uncovered by said cover plate, saidplatform cooling circuit includes an impingement cavity that is separatefrom said serpentine cavity, said cover plate is positioned to coversaid impingement cavity, the platform extends circumferentially betweenfirst and second mate faces, and the impingement cavity is positionedadjacent a trailing edge of the platform and extends from a positionadjacent the first mate face to a position adjacent the second mateface.
 17. The component as recited in claim 16, wherein the impingementcavity is generally L-shaped.